Design and Experimental Results for the S411 Airfoil

Transcription

1 RDECOM TR 10-D-111 U.S. ARMY RESEARCH, DEVELOPMENT AND ENGINEERING COMMAND TITLE: AUTHOR: COMPANY NAME: COMPANY ADDRESS: Design and Experimental Results for the S411 Airfoil Dan M. Somers and Mark D. Maughmer Airfoils, Incorporated 122 Rose Drive Port Matilda PA DATE: August 2010 FINAL REPORT: Contract Number W911W6-07-C-0047, SBIR Phase II, Topic Number A06-006, Proposal Number A DISTRIBUTION STATEMENT A Approved for public release; distribution is unlimited. Prepared for: U.S. ARMY RESEARCH, DEVELOPMENT AND ENGINEERING COMMAND, AVIATION APPLIED TECHNOLOGY DIRECTORATE, FORT EUSTIS, VA

2 AIRFOILS, INCORPORATED 122 ROSE DRIVE PORT MATILDA, PA USA WEBSITE WWW AIRFOILS.COM TELEPHONE (814) FACSIMILE (814) DESIGN AND EXPERIMENTAL RESULTS FOR THE S411 AIRFOIL DAN M. SOMERS AIRFOILS, INCORPORATED MARK D. MAUGHMER THE PENNSYLVANIA STATE UNIVERSITY AUGUST 2010

3 ABSTRACT A 14-percent-thick airfoil, the S411, intended for rotorcraft applications has been designed and analyzed theoretically and verified experimentally in The Pennsylvania State University Low-Speed, Low-Turbulence Wind Tunnel. The airfoil incorporates a 5-percentchord tab. The two primary objectives of high maximum lift and low profile drag have been achieved. The constraint on the pitching moment has been exceeded; that on the airfoil thickness, satisfied. The airfoil exhibits a docile stall. Comparisons of the theoretical and experimental results generally show good agreement. Comparisons with the S406 airfoil confirm the achievement of the objectives. INTRODUCTION Almost all airfoils in use on rotorcraft today were developed under the assumption that extensive laminar flow is not likely on a rotor. (See ref. 1, for example.) For the present application, however, given the relatively low Reynolds numbers and the precision blade manufacturing technique being employed, the achievement of laminar flow warrants exploration. The airfoil designed under the present effort is intended for the rotor of a small helicopter. To complement the design effort, an investigation was conducted in The Pennsylvania State University Low-Speed, Low-Turbulence Wind Tunnel (ref. 2) to obtain the basic, lowspeed, two-dimensional aerodynamic characteristics of the airfoil. The results have been compared with predictions from the method of references 3 and 4 and from the method of reference 5. The results have also been compared with those for the S406 airfoil (ref. 6), which has similar design specifications. SYMBOLS Values are given in both SI and U.S. Customary Units. Measurements and calculations were made in U.S. Customary Units. p l p pressure coefficient, C p c airfoil chord, mm (in.) q c c section chord-force coefficient, C p d z c - c d h section profile-drag coefficient, c d ' d --, except post stall, c sinα + c c cosα c n c d ' point drag coefficient (ref. 7) Wake c l section lift coefficient, c n cosα c d tanα

4 c m c n h M section pitching-moment coefficient about quarter-chord point, C x p x d -- c c z C p - c z d - + c section normal-force coefficient, horizontal width in wake profile, mm (in.) free-stream Mach number C p d x -- c p static pressure, Pa (lbf/ft 2 ) q dynamic pressure, Pa (lbf/ft 2 ) R t x y z α Reynolds number based on free-stream conditions and airfoil chord airfoil thickness, mm (in.) airfoil abscissa, mm (in.) model span station, y = 0 at midspan, mm (in.) airfoil ordinate, mm (in.) angle of attack relative to x-axis, deg Subscripts: l ll max min S T ul local point on airfoil lower limit of low-drag range maximum minimum separation transition upper limit of low-drag range 0 zero lift free-stream conditions 2

5 Abbreviations: L. lower surface S. boundary-layer separation location, x S /c T. boundary-layer transition location, x T /c U. upper surface AIRFOIL DESIGN OBJECTIVES AND CONSTRAINTS The airfoil design specifications are contained in table I. Two primary objectives are evident. The first objective is to achieve a maximum lift coefficient of 1.25 at a Mach number of 0.30 and a Reynolds number of and a maximum lift coefficient of 1.20 at a Mach number of 0.40 and a Reynolds number of A requirement related to this objective is that the maximum lift coefficient not decrease significantly with transition fixed near the leading edge on both surfaces. In addition, the airfoil should exhibit docile stall characteristics. The second objective is to obtain low profile-drag coefficients from a lift coefficient of 0.10 at a Mach number of 0.70 and a Reynolds number of to a lift coefficient of 0.65 at a Mach number of 0.45 and a Reynolds number of Three major constraints were placed on the design of the airfoil. First, the zero-lift pitching-moment coefficient must be 0 ± at a Mach number of 0.75 and a Reynolds number of with transition fixed at 10-percent chord on the upper and lower surfaces and 0 ± at a Mach number of 0.45 and a Reynolds number of with transition free. Second, the airfoil must incorporate a tab having a length of 5-percent chord and a thickness of percent chord. Third, the airfoil thickness must equal 14-percent chord with the tab. The specifications for this airfoil are similar to those for the S406 airfoil (ref. 6), except the pitching-moment constraint is tighter. 3

6 PHILOSOPHY Given the above objectives and constraints, certain characteristics of the design are apparent. The following sketch illustrates a drag polar that meets the goals for this design C c l.65 B.10 0 A c d Sketch 1 The desired airfoil shape can be traced to the pressure distributions that occur at the various points in sketch 1. Point A is the lower limit of the low-drag range of lift coefficients; point B, the upper limit. The profile-drag coefficient at point B is not as low as at point A, unlike the polars of many laminar-flow airfoils where the drag coefficient within the laminar bucket is nearly constant. (See, for example, ref. 8.) This characteristic is related to the elimination of significant (i.e., drag-producing) laminar separation bubbles on the upper surface for the design range of Reynolds numbers. (See ref. 9.) The drag coefficient increases rapidly outside the low-drag, lift-coefficient range because boundary-layer transition moves quickly toward the leading edge with increasing (or decreasing) lift coefficient. This feature results in a leading edge that produces a suction peak at higher lift coefficients, which ensures that transition on the upper surface will occur very near the leading edge. Thus, the maximum lift coefficient, point C, occurs with turbulent flow along the entire upper surface and, therefore, should be relatively insensitive to roughness at the leading edge. An unusual design approach was taken for this airfoil. Rather than design a thicker airfoil and then add the required tab, the airfoil was designed from the outset for the specified thickness including the tab. Specifically, the airfoil was initially designed with a trailing-edge shape that geometrically and aerodynamically approximated the tab. This shape was then modified to the required tab geometry. Accordingly, the performance of the final, tabbed airfoil is likely better than that of an airfoil altered by the addition of a relatively arbitrary tab. 4

7 From the preceding discussion, the pressure distributions along the polar can be deduced. The pressure distribution at point A for the airfoil shape with the pseudo tab should look something like sketch 2. Sketch 2 To achieve low drag, a favorable pressure gradient is desirable along the upper surface to about 25-percent chord. Aft of this point, a short region having a shallow, adverse pressure gradient (i.e., a transition ramp ) promotes the efficient transition from laminar to turbulent flow (ref. 10). The transition ramp is followed by a very slightly convex pressure recovery. The specific pressure recovery employed represents a compromise between maximum lift, drag, pitching moment, stall characteristics, and drag divergence. The steeper, adverse pressure gradient aft of about 90-percent chord is a separation ramp, originally proposed by F. X. Wortmann, 1 which confines turbulent separation to a small region near the trailing edge. By constraining the movement of the separation point at high angles of attack, higher lift coefficients can be achieved with little drag penalty. This feature has the added benefit of promoting docile stall characteristics. (See ref. 11.) Along the lower surface, the pressure gradient is initially adverse, then zero, and then favorable to about 70-percent chord. Thus, transition is imminent over the entire forward portion of the lower surface. (See ref. 12.) This concept allows a wide low-drag range to be achieved and increases the loading in the leading-edge region. The forward loading serves to balance, with respect to the pitching-moment constraint, the aft loading, both of which contribute to the achievement of the specified maximum lift coefficient and low profile-drag coefficients. This region is followed by a transition ramp and then a roughly linear pressure recovery. The pressure recovery must begin farther forward than optimum for low drag and 1 Director, Institute for Aerodynamics and Gas Dynamics, University of Stuttgart, Germany,

8 the constrained pitching moment to alleviate separation at lower lift coefficients, especially with transition fixed near the leading edge. The amounts of pressure recovery on the upper and lower surfaces are determined by the airfoil-thickness and pitching-moment constraints. At point B, the pressure distribution should look like sketch 3. Sketch 3 No suction peak exists at the leading edge. Instead, a rounded peak occurs aft of the leading edge, which allows some laminar flow, although not to the extent of point A. EXECUTION Given the pressure distributions previously discussed, the design of the airfoil is reduced to the inverse problem of transforming the pressure distributions into an airfoil shape. The Eppler Airfoil Design and Analysis Code (refs. 3 and 4) was used because of its unique capability for multipoint design and because of confidence gained during the design, analysis, and experimental verification of many other airfoils. (See ref. 13, for example.) The code also offers useful options for the modification of the airfoil geometry with respect to the tab. The airfoil is designated the S411. The airfoil shape incorporates a tab that is 5-percent-chord long and percent-chord thick, which satisfies the design constraint. 6

9 The airfoil shape and coordinates are available from Airfoils, Incorporated. The airfoil thickness is percent chord, which satisfies the design constraint. THEORETICAL PROCEDURE The theoretical results are predicted using the method of references 3 and 4 (PROFIL07), commonly known as the Eppler code, and the method of reference 5 (MSES 3.0). Critical amplification factors of 11 and 9 were specified for the computations using the method of references 3 and 4 and the method of reference 5, respectively. Because the maximum lift coefficient computed by the method of references 3 and 4 is not always realistic, an empirical criterion has been applied to the computed results. The criterion assumes the maximum lift coefficient has been reached if the drag coefficient of the upper surface reaches a certain value that is a function of the Reynolds number and the wind-tunnel facility. It should also be noted that the compressibility correction (ref. 14) incorporated in the method of references 3 and 4 is invalid if the local flow is supersonic. Because the free-stream Mach number for all wind-tunnel test conditions did not exceed 0.2, the flow can be considered essentially incompressible for the purpose of comparing the theoretical and experimental results. This allows the fast, subcritical flow solver of the method of reference 5 to be used. EXPERIMENTAL PROCEDURE WIND TUNNEL The Pennsylvania State University Low-Speed, Low-Turbulence Wind Tunnel (ref. 2) is a closed-throat, single-return, atmospheric tunnel (fig. 1). The test section is cm (39.9 in.) high by cm (58.1 in.) wide (fig. 2). Electrically actuated turntables provide positioning and attachment for the two-dimensional model. The turntables are flush with the top and bottom tunnel walls and rotate with the model. The axis of rotation coincided approximately with the midchord of the model, which was mounted vertically between the turntables. The gaps between the model and the turntables were sealed. The turbulence intensity in the test section is approximately 0.05 percent at 46 m/s (150 ft/s). MODEL The aluminum, wind-tunnel model was fabricated by Skytop Aerospace, Bellefonte, Pennsylvania, using a numerically controlled milling machine. The model had a chord of mm (18.00 in.) and a span of cm (42.50 in.) and, thus, extended through both turntables. Upper- and lower-surface orifices were located to one side of midspan at the staggered positions listed in table II. All the orifices were 0.51 mm (0.020 in.) in diameter with their axes perpendicular to the surface. The surfaces of the model were sanded to ensure an 7

10 aerodynamically smooth finish. The measured model contour was within 0.13 mm (0.005 in.) of the prescribed shape. WAKE-SURVEY PROBE A total- and static-pressure, wake-survey probe (fig. 3) was mounted from the top tunnel wall (fig. 2). The probe was positioned at midspan and automatically aligned with the wake-centerline streamline. A traverse mechanism incrementally positioned the probe to survey the wake. The increment was 1.27 mm (0.050 in.) for traverses less than mm (10.00 in.) and 2.54 mm (0.100 in.) for longer traverses, which were occasionally required near the maximum lift coefficient. The tip of the probe was located 0.83 chord downstream of the trailing edge of the model. INSTRUMENTATION Basic tunnel pressures and the wake pressures were measured with precision transducers. Measurements of the pressures on the model were made by an automatic pressurescanning system utilizing precision transducers. Data were obtained and recorded by an electronic data-acquisition system. METHODS The pressures measured on the model were reduced to standard pressure coefficients and numerically integrated to obtain section normal-force and chord-force coefficients and section pitching-moment coefficients about the quarter-chord point. Section profile-drag coefficients were computed from the wake total and static pressures by the method of reference 7. Wake surveys were not performed, however, at most post-stall angles of attack, in which case, the profile-drag coefficients were computed from the normal- and chord-force coefficients. Standard, low-speed, wind-tunnel boundary corrections (ref. 15) have been applied to the data. It should be noted, however, that the pressure distributions themselves are uncorrected. The wake-survey-probe total-pressure-tube displacement correction (ref. 7) has been taken into account. TESTS The model was tested at Reynolds numbers based on airfoil chord of , , , and with transition free (smooth), with transition fixed by roughness at 10-percent chord on the upper and lower surfaces to simulate a possible manufacturing deficiency, and with transition fixed by roughness near the leading edge, 2-percent chord on the upper surface and 7-percent chord on the lower surface, to simulate full-chord, 8

11 turbulent flow. Using the method of reference 16, the grit roughness was sized for lower lift coefficients for the 10-percent-chord location and for higher lift coefficients for the locations near the leading edge. The grit was sparsely distributed along 3-mm (0.1-in.) wide strips applied to the model with lacquer. (See table III.) The Mach number did not exceed 0.2 for any test condition. It should be noted that the test Mach numbers are much lower than the operational values of the intended application. Starting from 0, the angle of attack was increased to post-stall values. The angle of attack was then decreased from 0 to below that for zero lift. DISCUSSION OF RESULTS THEORETICAL RESULTS Pressure Distributions The inviscid pressure distributions at various angles of attack at Mach numbers of 0.30 and 0.45 predicted using the method of references 3 and 4 are shown in figure 4. The (viscous) pressure distributions at various angles of attack at a Mach number of 0.70 and a Reynolds number of predicted using the method of reference 5 are shown in figure 5. Section Characteristics The section characteristics at three of the design conditions with transition free and transition fixed are shown in figures 6 through 8. Based on the predictions, all the design objectives and constraints have been met, except for the zero-lift pitching-moment coefficient, which exceeds the constraint. EXPERIMENTAL RESULTS Pressure Distributions The pressure distributions at various angles of attack for a Reynolds number of and a Mach number of 0.10 with transition free are shown in figure 9. At an angle of attack of 2.01 (fig. 9(a)), transition probably occurs around 60-percent chord on the upper surface and near the leading edge on the lower surface. At an angle of attack of 1.00 (fig. 9(a)), which corresponds approximately to the lower limit of the low-drag, lift-coefficient range, a short laminar separation bubble is evident on the lower surface around 85-percent chord. As the angle of attack is increased, a short laminar separation bubble becomes more evident on the upper surface and moves forward, whereas the bubble on the lower surface remains relatively fixed (figs. 9(a) 9(c)). At an angle of attack of 8.16 (fig. 9(c)), turbulent, 9

12 trailing-edge separation occurs on the upper surface. The amount of separation increases with increasing angle of attack (figs. 9(c) and 9(d)). The maximum lift coefficient occurs between the angles of attack of 12 and 13 (figs. 9(c) and 9(d)). As the angle of attack is increased further, the separation point continues to move forward, although the leading-edge pressure peak does not fall (fig. 9(d)). Section Characteristics The section characteristics with transition free and transition fixed are shown in figure 10 and tabulated in the appendix. For a Reynolds number of and a Mach number of 0.10 with transition free (fig. 10(c)), the maximum lift coefficient is For a Reynolds number of and a Mach number of 0.16 with transition free (fig. 10(d)), the lower limit of the low-drag range of lift coefficients is approximately 0.05, the maximum lift-to-drag ratio occurs at a lift coefficient of about 0.96, and the zero-lift pitching-moment coefficient is (Because the upper limit of the low-drag, lift-coefficient range is not sharply defined, a precise value for the upper limit cannot be given.) The effects of Reynolds number on the section characteristics are summarized in figure 11. In general, with transition free, the lift-curve slope, the maximum lift coefficient, and the lower limit of the low-drag range increase with increasing Reynolds number. The zero-lift angle of attack, the profile-drag coefficients, and the pitching-moment coefficients, including the zero-lift value, generally decrease with increasing Reynolds number. The airfoil exhibits docile stall characteristics that become less docile with increasing Reynolds number. The effect of fixing transition on the section characteristics is shown in figure 10. In general, the zero-lift angle of attack, the lift-curve slope, the maximum lift coefficient, and the pitching-moment coefficients, including the zero-lift value, decrease with transition fixed. These results are primarily a consequence of the boundary-layer displacement effect, which decambers the airfoil because the displacement thickness is greater with transition fixed than with transition free. In addition, the maximum lift coefficient decreases with transition fixed because the roughness induces earlier trailing-edge separation. Accordingly, the decrease in maximum lift coefficient is inversely proportional to the roughness location, averaging 1 percent over the test Reynolds number range with transition fixed at 10-percent chord on the upper and lower surfaces and 9 percent with transition fixed at 2-percent chord on the upper surface and 7-percent chord on the lower surface. The elimination of the laminar separation bubble on the lower surface (see fig. 9) by the roughness also contributes to the decrease in the pitching-moment coefficient. The drag coefficients are, of course, generally affected adversely by the roughness. The stall characteristics are less docile with transition fixed. It should be noted that, for most test conditions, the Reynolds number based on local velocity and boundary-layer displacement thickness at the forward roughness locations, 2-percent and 7-percent chord, is too low to support turbulent flow. (See ref. 17.) Accordingly, to force transition, the roughness must be so large that it increases the displacement thickness, which abnormally decreases the lift coefficient and the magnitude of the pitchingmoment coefficient and increases the drag coefficient. Conversely, at low lift coefficients, the 10

13 roughness on the upper surface, which is sized for higher lift coefficients, is too small to force transition, resulting in incorrectly low drag coefficients. For Reynolds numbers of and (figs. 10(a) and 10(b)), the roughness near the leading edge alleviates the laminar separation bubble on the lower surface, thereby actually decreasing the drag coefficient. The variations of maximum lift coefficient and minimum profile-drag coefficient with Reynolds number are shown in figures 12 and 13, respectively. With transition free, the maximum lift coefficient increases with increasing Reynolds number, whereas the minimum profile-drag coefficient decreases, which are typical trends for most airfoils. (The minimum drag coefficient with transition fixed at 2-percent chord on the upper surface and 7-percent chord on the lower surface is too low because the roughness is too small to force transition at lower lift coefficients, as previously discussed.) COMPARISON OF THEORETICAL AND EXPERIMENTAL RESULTS Pressure Distributions The comparison of the theoretical and experimental pressure distributions at various angles of attack is shown in figure 14. It should be noted that the pressure distributions predicted using the method of references 3 and 4 (PROFIL07) are inviscid and incompressible, whereas the pressure distributions predicted using the method of reference 5 (MSES 3.0) as well as the experimental pressure distributions were obtained for a Reynolds number of and a Mach number of 0.10 with transition free. It should also be noted that the theoretical lift coefficient from the method of references 3 and 4 is calculated from the liftcurve slope and the angle of attack relative to the zero-lift line, whereas the lift coefficient from the method of reference 5 and from the experiment is derived from the integrated pressure distribution. (See refs. 3 5 and 7.) Thus, at a given lift coefficient, the pressure distribution predicted using the method of references 3 and 4 does not necessarily have the same area as the measured pressure distribution. It should be noticed that the angle of attack shown in figure 14 is the theoretical value from the method of references 3 and 4, not the experimental value. Also, the lift coefficient shown in this figure only is the uncorrected value. With respect to the method of references 3 and 4, at a lift coefficient of 0.13 (fig. 14(a)), which is near the lower limit of the low-drag range, the pressure coefficients and the pressure gradients agree well, except where laminar separation bubbles are present and along the upper surface in the vicinity of the start of the tab. The latter disparity is probably the result of the displacement effect. At a lift coefficient of 0.61 (fig. 14(b)), although the pressure coefficients do not match exactly, the pressure gradients agree reasonably well, again except where bubbles are present and along the upper surface near the start of the tab. At a lift coefficient of 1.14 (fig. 14(c)), which is near the experimental maximum lift coefficient, the agreement is poor because the effect of the upper-surface, trailing-edge separation on the pressure distribution is not modelled in the method of references 3 and 4. With respect to the method of reference 5, at a lift coefficient of 0.13 (fig. 14(a)), the pressure coefficients and the pressure gradients agree remarkably well. The location of the 11

14 lower-surface laminar separation bubble is predicted well, but that of the upper-surface bubble is aft of the measured location. At a lift coefficient of 0.61 (fig. 14(b)), although the pressure coefficients do not match exactly, the pressure gradients agree well. The predicted location of the upper-surface bubble is again aft of the measured location. At a lift coefficient of 1.14 (fig. 14(c)), the agreement is less exact because the extent of the upper-surface, trailing-edge separation and, in turn, its effect on the overall circulation are underpredicted by the method of reference 5. Section Characteristics The comparison of the theoretical and experimental section characteristics with transition free is shown in figure 15. The previously discussed empirical criterion applied to the results from the method of references 3 and 4 (PROFIL07) underestimates the maximum lift coefficient by an average of 11 percent over the test Reynolds number range. The method of reference 5 (MSES 3.0) overpredicts the maximum lift coefficient by an average of 17 percent. The method of references 3 and 4 generally overpredicts the profile-drag coefficients, whereas the method of reference 5 generally underpredicts the drag coefficients. Both methods predict the zero-lift angle of attack, the lift-curve slope, the lower limit of the low-drag range, and the zero-lift pitching-moment coefficient reasonably well. Both methods also predict more positive pitching-moment coefficients and underpredict the effect of the trailingedge separation on the lift coefficient at higher angles of attack. Overall, the general agreement improves with increasing Reynolds number. The comparison of the theoretical and experimental section characteristics with transition fixed at 10-percent chord on the upper and lower surfaces is shown in figure 16. In general, the predicted characteristics show similar tendencies as with transition free. The method of references 3 and 4 underpredicts the magnitude of the pitching-moment coefficients at lower angles of attack, whereas the method of reference 5 predicts the values well. The comparison of the theoretical and experimental section characteristics with transition fixed at 2-percent chord on the upper surface and 7-percent chord on the lower surface is shown in figure 17. In general, the predicted characteristics show similar tendencies as with transition free, although the general agreement is poorer, particularly with respect to the drag coefficients, probably because of the abnormalities introduced by the roughness, as discussed previously. The empirical criterion applied to the results from the method of references 3 and 4 underpredicts the maximum lift coefficient by an average of 6 percent, whereas the method of reference 5 overpredicts the maximum lift coefficient by an average of 28 percent. The method of reference 5 overpredicts the magnitude of the pitching-moment coefficients at lower angles of attack, whereas the method of references 3 and 4 predicts the values well. Given the abrupt, contour changes introduced by the tab, the agreement between the theoretical and experimental section characteristics is remarkably good overall. 12

15 COMPARISON WITH S406 AIRFOIL The section characteristics of the S411 airfoil for a Reynolds number of and a Mach number of 0.10 with transition free are compared with those of the S406 airfoil, which has similar design specifications, in figure 18. The maximum lift coefficients and the profiledrag coefficients at a lift coefficient of 0.4 are compared in figures 19 and 20, respectively. The maximum lift coefficients with transition free are nearly identical, but the S411 airfoil suffers a larger effect of roughness. The S411 airfoil exhibits substantially higher drag coefficients but also substantially less negative pitching-moment coefficients. CONCLUDING REMARKS A percent-thick airfoil, the S411, intended for rotorcraft applications has been designed and analyzed theoretically and verified experimentally in The Pennsylvania State University Low-Speed, Low-Turbulence Wind Tunnel. The airfoil incorporates a 5-percentchord tab. The two primary objectives of a high maximum lift coefficient and low profiledrag coefficients have been achieved. The constraint on the zero-lift pitching-moment coefficient has been exceeded; the constraint on the airfoil thickness has been satisfied. The airfoil exhibits docile stall characteristics. Comparisons of the theoretical and experimental results generally show good agreement. Comparisons with the S406 airfoil confirm the achievement of the objectives. ACKNOWLEDGMENTS This effort was sponsored by the U.S. Army. Preston B. Martin served as the technical monitor. 13

16 REFERENCES 1. Noonan, Kevin W.: Aerodynamic Characteristics of Two Rotorcraft Airfoils Designed for Application to the Inboard Region of a Main Rotor Blade. NASA TP-3009, Brophy, Christopher M.: Turbulence Management and Flow Qualification of The Pennsylvania State University Low Turbulence, Low Speed, Closed Circuit Wind Tunnel. M. S. Thesis, Pennsylvania State Univ., Eppler, Richard: Airfoil Design and Data. Springer-Verlag (Berlin), Eppler, Richard: Airfoil Program System PROFIL07. User s Guide. Richard Eppler, c Drela, M.: Design and Optimization Method for Multi-Element Airfoils. AIAA Paper , Feb Somers, Dan M.; and Maughmer, Mark D.: Design and Experimental Results for the S406 Airfoil. U.S. Army RDECOM TR 10-D-107, (Available from DTIC.) 7. Pankhurst, R. C.; and Holder, D. W.: Wind-Tunnel Technique. Sir Isaac Pitman & Sons, Ltd. (London), Abbott, Ira H.; Von Doenhoff, Albert E.; and Stivers, Louis S., Jr.: Summary of Airfoil Data. NACA Rep. 824, (Supersedes NACA WR L-560.) 9. Eppler, Richard; and Somers, Dan M.: Airfoil Design for Reynolds Numbers Between 50,000 and 500,000. Proceedings of the Conference on Low Reynolds Number Airfoil Aerodynamics, UNDAS-CP-77B123, Univ. of Notre Dame, June 1985, pp Wortmann, F. X.: Experimental Investigations on New Laminar Profiles for Gliders and Helicopters. TIL/T.4906, British Minist. Aviat., Mar (Translated from Z. Flugwissenschaften, Bd. 5, Heft 8, Aug. 1957, S ) 11. Maughmer, Mark D.; and Somers, Dan M.: Design and Experimental Results for a High- Altitude, Long-Endurance Airfoil. J. Aircr., vol. 26, no. 2, Feb. 1989, pp Eppler, R.: Laminar Airfoils for Reynolds Numbers Greater Than B , Apr (Available from NTIS as N ; translated from Ingenieur-Archiv, Bd. 38, Heft 4/5, 1969, S ) 13. Somers, Dan M.: Subsonic Natural-Laminar-Flow Airfoils. Natural Laminar Flow and Laminar Flow Control, R. W. Barnwell and M. Y. Hussaini, eds., Springer-Verlag New York, Inc., 1992, pp

17 14. Labrujere, Th. E.; Loeve, W.; and Sloof, J. W.: An Approximate Method for the Determination of the Pressure Distribution on Wings in the Lower Critical Speed Range. Transonic Aerodynamics. AGARD CP No. 35, Sept. 1968, pp Allen, H. Julian; and Vincenti, Walter G.: Wall Interference in a Two-Dimensional-Flow Wind Tunnel, With Consideration of the Effect of Compressibility. NACA Rep. 782, (Supersedes NACA WR A-63.) 16. Braslow, Albert L.; and Knox, Eugene C.: Simplified Method for Determination of Critical Height of Distributed Roughness Particles for Boundary-Layer Transition at Mach Numbers From 0 to 5. NACA TN 4363, Schubauer, G. B.; and Klebanoff, P. S.: Contributions on the Mechanics of Boundary- Layer Transition. NACA Rep. 1289,

18 TABLE I.- AIRFOIL DESIGN SPECIFICATIONS Parameter Objective/ Constraint Mach Number M Reynolds Number R Priority Minimum lift coefficient c l,min Low Maximum lift coefficient 1.25 c l,max High Lower limit of low-drag, lift-coefficient range c l,ll Medium Upper limit of low-drag, lift-coefficient range c l,ul Medium Zero-lift pitching-moment 0 ± coefficient c m,0 0 ± High Thickness t/c 0.14 with tab Medium Other requirements: Maximum lift coefficient c l,max independent of leading-edge roughness Docile stall characteristics 5-percent-chord tab with thickness of percent chord 1 With transition fixed at 10-percent chord on upper and lower surfaces. 2 With transition free. 16

19 TABLE II.- MODEL ORIFICE LOCATIONS [c = mm (18.00 in.)] Upper Surface Lower Surface x/c y, mm (in.) x/c y, mm (in.) ( 5.10) ( 6.33) ( 5.06) ( 6.29) ( 5.02) ( 6.23) ( 4.98) ( 6.16) ( 4.93) ( 6.07) ( 4.87) ( 5.97) ( 4.79) ( 5.85) ( 4.70) ( 5.72) ( 4.60) ( 5.58) ( 4.49) ( 5.43) ( 4.37) ( 5.27) ( 4.24) ( 5.11) ( 4.10) ( 4.94) ( 3.95) ( 4.77) ( 3.80) ( 4.60) ( 3.64) ( 4.43) ( 3.48) ( 4.26) ( 3.32) ( 4.10) ( 3.16) ( 3.95) ( 3.00) ( 3.81) ( 3.00) ( 3.69) ( 3.00) ( 3.58) ( 3.00) ( 3.47) ( 3.00) ( 3.37) ( 3.00) ( 3.28) ( 3.00) ( 3.20) ( 3.00) ( 3.14) ( 3.00) ( 3.14) ( 2.06) ( 3.14) ( 2.00) ( 3.26) ( 1.94) ( 1.88) 17

20 TABLE III.- ROUGHNESS LOCATIONS AND SIZES Upper surface Lower surface R x/c Grit number Nominal size, mm (in.) x/c Grit number Nominal size, mm (in.) (0.0280) (0.0232) (0.0117) (0.0280) (0.0232) (0.0232) (0.0083) (0.0232) (0.0165) (0.0138) (0.0059) (0.0138) (0.0117) (0.0098) (0.0049) (0.0098) 18

21 19 Figure 1.- The Pennsylvania State University Low-Speed, Low-Turbulence Wind Tunnel.

22 20 Figure 2.- S411 airfoil model and wake-survey probe mounted in test section.

23 Static-pressure connection Total-pressure connection 6.4 mm (0.25 in.) 57.2 mm (2.25 in.) 25.4 mm (1.00 in.) 1.60 mm (0.063 in.) 5 equally spaced orifices, 0.64-mm (0.025-in.) diameter Figure 3.- Wake-survey probe. 21

24 (a) M = Figure 4.- Theoretical (inviscid) pressure distributions. 22

25 (b) M = Figure 4.- Concluded. 23

26 Figure 5.- Theoretical pressure distributions at M = 0.70 and R = with transition free. 24

27 (a) Transition free. 25 Figure 6.- Theoretical section characteristics at M = 0.30 and R =

28 26 (b) Transition fixed. Figure 6.- Concluded.

29 (a) Transition free. 27 Figure 7.- Theoretical section characteristics at M = 0.45 and R =

30 28 (b) Transition fixed. Figure 7.- Concluded.

31 (a) Transition free. 29 Figure 8.- Theoretical section characteristics at M = 0.70 and R =

32 30 (b) Transition fixed. Figure 8.- Concluded.

33 (a) α = 2.01, 1.00, 0.02, 1.04, and Figure 9.- Experimental pressure distributions for R = and M = 0.10 with transition free. 31

34 (b) α = 3.08, 4.10, 5.11, 6.13, and Figure 9.- Continued. 32

35 (c) α = 8.16, 9.18, 10.19, 11.20, and Figure 9.- Continued. 33

36 (d) α = 13.21, 14.20, 15.19, 16.18, 17.17, and Figure 9.- Concluded. 34

37 (a) R = and M = Figure 10.- Experimental section characteristics with transition free and transition fixed.

38 36 (b) R = and M = Figure 10.- Continued.

39 (c) R = and M = Figure 10.- Continued.

40 38 (d) R = and M = Figure 10.- Concluded.

41 (a) Transition free. 39 Figure 11.- Effects of Reynolds number on experimental section characteristics.

42 40 (b) Transition fixed at 0.10c U. and 0.10c L. Figure 11.- Continued.

43 (c) Transition fixed at 0.02c U. and 0.07c L. 41 Figure 11.- Concluded.

44 42 Figure 12.- Variation of experimental maximum lift coefficient with Reynolds number.

45 Figure 13.- Variation of experimental minimum profile-drag coefficient with Reynolds number. 43

46 (a) c l = Figure 14.- Comparison of theoretical and experimental pressure distributions. 44

47 (b) c l = Figure 14.- Continued. 45

48 (c) c l = Figure 14.- Concluded. 46

49 (a) R = Figure 15.- Comparison of theoretical and experimental section characteristics with transition free.

50 48 (b) R = Figure 15.- Continued.

51 (c) R = Figure 15.- Continued.

52 50 (d) R = Figure 15.- Concluded.

53 (a) R = Figure 16.- Comparison of theoretical and experimental section characteristics with transition fixed at 0.10c U. and 0.10c L.

54 52 (b) R = Figure 16.- Continued.

55 (c) R = Figure 16.- Continued.

56 54 (d) R = Figure 16.- Concluded.

57 (a) R = Figure 17.- Comparison of theoretical and experimental section characteristics with transition fixed at 0.02c U. and 0.07c L.

58 56 (b) R = Figure 17.- Continued.

59 (c) R = Figure 17.- Continued.

60 58 (d) R = Figure 17.- Concluded.

61 59 Figure 18.- Comparison of section characteristics of S411 and S406 airfoils for R = and M = 0.1 with transition free.

62 Figure 19.- Comparison of maximum lift coefficients of S411 and S406 airfoils. Open symbols represent data with transition free; solid symbols, data with transition fixed at 0.02c U. and 0.07c L. Figure 20.- Comparison of drag coefficients at c l = 0.4 of S411 and S406 airfoils with transition free. 60

63 APPENDIX EXPERIMENTAL SECTION CHARACTERISTICS 61

64 R = , M = 0.05, transition free α, deg c l c d c m

65 R = , M = 0.05, transition fixed at 0.10c U. and 0.10c L. α, deg c l c d c m

66 R = , M = 0.05, transition fixed at 0.02c U. and 0.07c L. α, deg c l c d c m

67 R = , M = 0.07, transition free α, deg c l c d c m

68 R = , M = 0.07, transition fixed at 0.10c U. and 0.10c L. α, deg c l c d c m

69 R = , M = 0.07, transition fixed at 0.02c U. and 0.07c L. α, deg c l c d c m

70 R = , M = 0.10, transition free α, deg c l c d c m

71 R = , M = 0.10, transition fixed at 0.10c U. and 0.10c L. α, deg c l c d c m

72 R = , M = 0.10, transition fixed at 0.02c U. and 0.07c L. α, deg c l c d c m

73 R = , M = 0.16, transition free α, deg c l c d c m

74 R = , M = 0.16, transition fixed at 0.10c U. and 0.10c L. α, deg c l c d c m

75 R = , M = 0.16, transition fixed at 0.02c U. and 0.07c L. α, deg c l c d c m

76 REPORT DOCUMENTATION PAGE Form Approved OMB No Public reporting burden for this collection of information is estimated to average 1 hour per response, including the time for reviewing instructions, searching existing data sources, gathering and maintaining the data needed, and completing and reviewing this collection of information. Send comments regarding this burden estimate or any other aspect of this collection of information, including suggestions for reducing this burden to Department of Defense, Washington Headquarters Services, Directorate for Information Operations and Reports ( ), 1215 Jefferson Davis Highway, Suite 1204, Arlington, VA Respondents should be aware that notwithstanding any other provision of law, no person shall be subject to any penalty for failing to comply with a collection of information if it does not display a currently valid OMB control number. PLEASE DO NOT RETURN YOUR FORM TO THE ABOVE ADDRESS. 1. REPORT DATE (DD-MM-YYYY) xx REPORT TYPE FINAL REPORT 3. DATES COVERED (From - To) Sep 2007 Jun TITLE AND SUBTITLE 5a. CONTRACT NUMBER W911W6 07 C 0047 Design and Experimental Results for the S411 Airfoil 5b. GRANT NUMBER 5c. PROGRAM ELEMENT NUMBER 6. AUTHOR(S) 5d. PROJECT NUMBER Somers, Dan M. and Maughmer, Mark D. 5e. TASK NUMBER 5f. WORK UNIT NUMBER 7. PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES) 8. PERFORMING ORGANIZATION REPORT NUMBER Airfoils, Incorporated Attn: Dan M. Somers 122 Rose Drive Port Matilda PA SBIR Topic Number A Proposal Number A SPONSORING / MONITORING AGENCY NAME(S) AND ADDRESS(ES) 10. SPONSOR/MONITOR S ACRONYM(S) US Army Aviation Research, Development and Engineering Command (RDECOM) 11. SPONSOR/MONITOR S REPORT Aviation Applied Technology Directorate (AATD) NUMBER(S) Fort Eustis VA RDECOM TR 10 D DISTRIBUTION / AVAILABILITY STATEMENT Approved for public release; distribution is unlimited. 13. SUPPLEMENTARY NOTES UL Note: No proprietary / limited information may be included in the abstract. 14. ABSTRACT A 14 percent thick airfoil, the S411, intended for rotorcraft applications has been designed and analyzed theoretically and verified experimentally in The Pennsylvania State University Low Speed, Low Turbulence Wind Tunnel. The airfoil incorporates a 5 percent chord tab. The two primary objectives of high maximum lift and low profile drag have been achieved. The constraint on the pitching moment has been exceeded; that on the airfoil thickness, satisfied. The airfoil exhibits a docile stall. Comparisons of the theoretical and experimental results generally show good agreement. Comparisons with the S406 airfoil confirm the achievement of the objectives. 15. SUBJECT TERMS Airfoils, rotorcraft, laminar flow, wind tunnel 16. SECURITY CLASSIFICATION OF: 17. LIMITATION OF ABSTRACT a. REPORT unclassified b. ABSTRACT unclassified 18. NUMBER OF PAGES c. THIS PAGE unclassified UU 73 19a. NAME OF RESPONSIBLE PERSON Dan M. Somers 19b. TELEPHONE NUMBER (include area code) (814) Standard Form 298 (Rev. 8-98) Prescribed by ANSI Std. Z39.18